Embedded ceramic matrix composite turbine outer ring for aero-engine

Abstract

The invention discloses an embedded ceramic matrix composite turbine outer ring for an aero-engine in the technical field of aero-engine parts, and the embedded ceramic matrix composite turbine outer ring for the aero-engine comprises a turbine case, a plurality of turbine outer rings are arranged in the turbine case, the plurality of turbine outer rings are circumferentially distributed, and the turbine outer rings are arranged in the turbine case. A special-shaped bolt is arranged between the turbine outer ring and the turbine case, and the special-shaped bolt and the turbine outer ring are correspondingly arranged. By arranging the special-shaped bolt and spring composite connection structure, the problem that the connection mode is poor in connection reliability in application of the ceramic matrix composite is solved, firstly, the special-shaped bolt is matched with the limiting groove, and the pre-compression spring is combined to form elastic constraint; the continuous elastic force of the spring counteracts the thermal deformation stress caused by the difference of linear expansion coefficients of the CMC and the high-temperature alloy, meanwhile, working vibration is restrained, rigid collision between the turbine outer ring and the casing is avoided, and the blade tip gap fluctuation risk is remarkably reduced.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engine parts, and in particular to an embedded ceramic-based composite material turbine outer ring for an aero-engine. Background Art

[0002] As a major component of an aircraft engine, the performance of the turbine directly impacts engine efficiency. The turbine outer ring, a crucial component of the turbine, sits between the turbine casing and the turbine blades. It is a typical hot-end stator component, responsible for reducing tip clearance leakage and protecting the engine casing from erosion by high-temperature leaking gases.

[0003] Continuous fiber-reinforced ceramic matrix composites (CMCs) incorporate continuous fibers into the ceramic matrix to improve the inherent brittleness of the ceramic material, imparting reliable structural load-bearing capacity. Commonly used CMCs (CMCs) in aircraft engines include SiCf / SiC and Al2O3 / Al2O3. SiCf / SiC is the most widely used and holds the greatest promise for future applications. Its characteristics include high-temperature resistance (and high-temperature oxidation resistance), lightweight (low density, high specific strength), excellent stealth (for infrared and radar), and high strength and toughness (compared to conventional ceramics). This makes it a promising material for the hot-end components of aircraft gas turbine engines.

[0004] At present, the outer ring components of the turbine are mainly made of high-temperature alloy materials. However, the strength and stiffness of high-temperature alloy materials will drop significantly under high-temperature environments, which in turn affects the upper limit of the operating temperature of the outer ring components of the turbine. It is difficult to meet the high-temperature creep strength and high-temperature oxidation and corrosion resistance required for the safe and reliable operation of the engine, as well as the needs of reducing the weight of the engine structure and improving the power-to-weight ratio. It has become a bottleneck and restrictive factor for high thrust-to-weight ratio engines to achieve technical indicators.

[0005] Using ceramic-based composite materials (CMC) to replace high-temperature alloy materials to prepare aircraft engine turbine outer ring components can fully utilize CMC's characteristics of low density, excellent high-temperature mechanical properties and good thermal stability. It can improve the temperature resistance of parts by about 300°C, solve the local over-temperature and ablation deformation problems of existing metal turbine outer rings, reduce the amount of cooling air drawn from the compressor (no need to use cooling air); reduce the weight of the structure by 50% to 70%, and is also conducive to controlling the gap and reducing the diffusion of gas heat to the casing, improving turbine efficiency and reducing engine fuel consumption.

[0006] However, currently, turbine outer rings are often connected using bolts or hooks. Because the strength and stiffness of ceramic matrix composites are inferior to those of high-temperature alloys, bolting alone is unreliable. Furthermore, the significant difference in linear expansion coefficients between the SiCf / SiC composite and the high-temperature alloy connector causes vibration of the turbine outer ring on the turbine casing, affecting blade tip clearance and potentially causing serious failures such as blade rubbing and damage during high-speed turbine rotor rotation. Therefore, we propose a flush-mounted ceramic matrix composite turbine outer ring for aircraft engines.

[0007] Practical content In response to the shortcomings of the existing technology, the present invention provides an embedded ceramic-based composite turbine outer ring for aircraft engines. A composite connection structure is constructed by special-shaped bolts and springs to form an elastic constraint, which is used to compensate for thermal deformation differences and suppress vibration. The outer ring gap and spring buffer form an adaptive mechanism to avoid stress concentration, solving the problem of traditional bolt connections being easy to loosen and vibrate greatly in CMC material applications, resulting in uncontrolled blade tip clearance.

[0008] The objective of the present invention is achieved as follows: An embedded ceramic-based composite turbine outer ring for an aircraft engine comprises a turbine casing, a turbine outer ring is arranged inside the turbine casing, the number of the turbine outer rings is multiple, and the multiple turbine outer rings are distributed circumferentially, special-shaped bolts are arranged between the turbine outer ring and the turbine casing, the special-shaped bolts are arranged corresponding to the turbine outer ring, the special-shaped bolts pass through the turbine casing, a spring is arranged inside the turbine outer ring, a hollow area is formed on the turbine casing, a turbine rotor is arranged on the turbine casing, the turbine rotor is located in the hollow area, and a tip clearance is arranged between the turbine rotor and the turbine casing.

[0009] Optionally, an inner annular surface is provided on the turbine casing, an annular groove is formed on the inner annular surface, and the annular groove is provided corresponding to the outer ring of the turbine.

[0010] Optionally, a positioning hole is opened on the turbine casing, and the positioning hole is located in the annular groove. There are multiple positioning holes, and the multiple positioning holes are distributed in a circle.

[0011] Optionally, an outer ring gap is provided between adjacent turbine outer rings, and the outer ring gap is used to provide expansion space.

[0012] Optionally, positioning plates are provided on both sides of the turbine outer ring, and a positioning block is provided on the turbine outer ring, and the positioning block and the positioning plate are arranged in an I-shape.

[0013] Optionally, the positioning block is provided with two limiting slots, which are symmetrically distributed.

[0014] Optionally, both sides of the positioning groove are open, and inner walls on both sides of the positioning groove are provided with inner groove inclined surfaces, and the inner groove inclined surfaces on both sides are symmetrical.

[0015] Optionally, the special-shaped bolt includes a bolt column, a special-shaped head is fixedly connected to the bolt column, a bevel section is provided on the side of the special-shaped head, the bevel section is provided corresponding to the bevel in the groove, and the special-shaped bolt is provided corresponding to the limiting groove.

[0016] Optionally, one end of the spring contacts the limiting groove, and the other end of the spring contacts the special-shaped head, and the spring is in a compressed state.

[0017] Optionally, a gasket is sleeved on the outer side of the bolt column, the gasket is in contact with the outer side of the turbine casing, and a self-locking nut is threadedly connected to the bolt column.

[0018] Compared with the prior art, the present invention has the following beneficial effects: 1. A composite connection structure of special-shaped bolts and springs is proposed to address the problem of poor connection reliability in the application of ceramic-based composite materials. First, the special-shaped bolts are combined with limit grooves and pre-compressed springs to form elastic constraints. The continuous elastic force of the spring offsets the thermal deformation stress caused by the difference in linear expansion coefficient between CMC and high-temperature alloys, while suppressing operating vibrations, avoiding rigid collisions between the turbine outer ring and the casing, and significantly reducing the risk of tip clearance fluctuations.

[0019] 2. The linear expansion coefficients between ceramic-based composite materials and high-temperature alloy connectors are very different. Due to temperature changes, large thermal stress will be generated at the interface between the two materials. The problem of thermal expansion mismatch between ceramic-based composite materials and metal materials is solved by utilizing the scaling of the spring, ensuring that the turbine outer ring block and the turbine casing always maintain a close fit under any operating condition of the engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are merely embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the provided drawings without paying any creative work.

[0021] Figure 1 It is a schematic diagram of the overall structure provided by the present invention; Figure 2 The present invention provides Figure 1 A magnified view of part A; Figure 3 This is a schematic diagram of the turbine casing structure provided by the present invention; Figure 4This is a schematic diagram of the turbine outer ring structure provided by the present invention; Figure 5 This is a schematic diagram of the structure of the special-shaped bolt provided by the present invention; Figure 6 Schematic diagram of the spring structure provided by the present invention; Figure 7 It is a schematic diagram of the turbine rotor structure provided by the present invention.

[0022] In the figure: 1. Turbine outer ring; 11. Positioning plate; 12. Positioning block; 13. Limiting groove; 14. Inner inclined surface of groove; 15. Outer ring gap; 2. Turbine casing; 21. Inner ring surface; 22. Annular groove; 23. Positioning hole; 3. Special-shaped bolt; 31. Bolt column; 32. Special-shaped head; 33. Inclined section; 4. Spring; 5. Gasket; 6. Self-locking nut; 7. Turbine rotor; 8. Tip clearance. DETAILED DESCRIPTION

[0023] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0024] like Figures 1 to 7 The shown embodiment shows an embedded ceramic-based composite turbine outer ring for an aircraft engine, comprising a turbine casing 2, a turbine outer ring 1 being arranged inside the turbine casing 2, a plurality of turbine outer rings 1 being arranged, and the plurality of turbine outer rings 1 being distributed circumferentially, a special-shaped bolt 3 being arranged between the turbine outer ring 1 and the turbine casing 2, the special-shaped bolt 3 being arranged corresponding to the turbine outer ring 1, the special-shaped bolt 3 passing through the turbine casing 2, a spring 4 being arranged inside the turbine outer ring 1, a hollow area being formed on the turbine casing 2, a turbine rotor 7 being arranged on the turbine casing 2, the turbine rotor 7 being located in the hollow area, and a blade tip clearance 8 being arranged between the turbine rotor 7 and the turbine casing 2.

[0025] It should be noted that the embodiments provided by the present invention mainly solve the following problems: first, the turbine outer ring 1 is made of a ceramic matrix composite material. Due to the high melting point and electrical insulation properties of ceramic matrix composite materials, which are different from metals, it is impossible to connect them using traditional welding methods. At the same time, the brittleness of ceramic matrix composite materials also makes it impossible to connect them using traditional bolts. Therefore, when designing the flame tube, it is necessary to solve the problem of fixing the composite outer ring to the metal turbine casing 2; Secondly, due to the significant difference in linear expansion coefficients between the SiCf / SiC composite and the high-temperature alloy connector to which it is connected, the ceramic-based composite's thermal expansion coefficient at high temperatures is only approximately 1 / 3 of that of the metal. During engine operation, temperature fluctuations can generate significant residual thermal stresses at the interface between the two materials, severely impacting the composite's load-bearing capacity. The turbine outer ring 1 is located on the turbine casing 2, corresponding to the blade tips of the aircraft engine's high-pressure turbine rotor 7. During operation, it comes into direct contact with high-temperature combustion gases. Therefore, when connecting ceramic-based composite components to metal components, the thermal expansion mismatch between the two materials must be addressed.

[0026] Finally, there is the problem of controlling the tip clearance 8. The turbine outer ring 1 is mainly subjected to the erosion of high-temperature combustion gas during service, and needs to withstand certain mechanical excitation forces and aerodynamic loads. In order to pursue the smallest possible tip clearance 8 under stable working conditions, and avoid intermittent scraping between the blade tip and the outer ring during the engine transition state, or even damage the high-speed rotating turbine blades and cause major failures, it is necessary to ensure the centering positioning of the ceramic-based turbine outer ring 1 in the design, and fit tightly to the turbine casing 2, so as to reduce the vibration of the turbine outer ring 1 under mechanical external forces and aerodynamic loads.

[0027] Furthermore, the problem of ceramic-based materials being unable to be welded and traditional bolt connections being prone to brittle cracking is solved by using a special-shaped bolt 3 connection structure. The elastic element design effectively alleviates the thermal expansion difference between ceramics and metals. At the same time, the innovative positioning structure and elastic support system achieve precise control of the blade tip clearance 8, which not only ensures structural reliability under high temperatures, but also maintains the blade tip clearance 8 within the optimal range, significantly improving turbine efficiency and engine performance.

[0028] Specifically, the turbine casing 2 is provided with an inner annular surface 21 , and an annular groove 22 is formed on the inner annular surface 21 . The annular groove 22 is provided corresponding to the turbine outer ring 1 .

[0029] Furthermore, the annular groove 22 structure provides a precise radial positioning reference for the turbine outer ring 1, ensuring that the tip clearance 8 is controlled within the optimized range, thereby improving the turbine efficiency; secondly, the annular groove 22 structure allows the turbine outer ring 1 to freely expand along the annular groove 22 under high temperature conditions, thereby alleviating the thermal stress between the ceramic-based composite material and the metal casing and reducing the peak interface stress; thirdly, the annular groove 22 and the special-shaped bolt 3 form a double constraint.

[0030] Specifically, the turbine casing 2 is provided with positioning holes 23 , which are located in the annular groove 22 . There are multiple positioning holes 23 , which are distributed in a circular pattern.

[0031] Furthermore, the ceramic matrix composite turbine outer ring 1 is evenly distributed circumferentially on the inner side of the turbine casing 2 to reduce tip clearance leakage and protect the engine casing from erosion by high-temperature leaking gas. An array of locating holes 23 provides precise circumferential positioning for each outer ring segment, ensuring uniform tip clearance 8 and reducing leakage losses. Secondly, the multiple independently arranged ceramic outer rings form a floating seal structure that adapts to thermal deformation under high-temperature conditions. Thirdly, the locating holes 23 and the special-shaped bolts 3 form a dual constraint, allowing for compensation for radial thermal expansion. Specifically, an outer ring gap 15 is provided between adjacent turbine outer rings 1 , and the outer ring gap 15 is used to provide expansion space.

[0032] Furthermore, an outer ring gap 15 is provided around each turbine outer ring 1 to provide expansion space around the turbine outer ring 1 in a hot state to prevent it from being squeezed, deformed or damaged. First, the precise expansion space allows the ceramic-based composite turbine outer ring 1 to expand freely under high-temperature conditions, avoiding the problem of thermal stress concentration, so that the turbine outer ring 1 can still maintain a certain dynamic sealing gap after high-temperature expansion, thereby improving the tip efficiency. In addition, the outer ring gap 15 allows a single turbine outer ring 1 to deform independently, avoiding brittle fracture of the ceramic material caused by overall extrusion and extending its service life.

[0033] Specifically, positioning plates 11 are provided on both sides of the turbine outer ring 1 , and positioning blocks 12 are provided on the turbine outer ring 1 . The positioning blocks 12 and the positioning plates 11 are arranged in an I-shape.

[0034] Furthermore, the positioning block 12 on the turbine outer ring 1 is arranged corresponding to the annular groove 22. During installation, the positioning block 12 is inserted into the annular groove 22. The positioning plates 11 on both sides of the turbine outer ring 1 fit tightly with the turbine casing 2, providing precise radial positioning references to ensure the uniformity and stability of the tip clearance 8. The I-shaped design forms a two-way constraint, which can effectively resist the centrifugal force generated by high-speed rotation and allow necessary thermal expansion deformation. The close fit between the positioning plate 11 and the casing enhances the overall stiffness of the structure and significantly suppresses vibration and deformation. At the same time, the modular positioning method simplifies the assembly process and facilitates maintenance and replacement.

[0035] Specifically, a limiting groove 13 is provided on the positioning block 12, and the number of the limiting grooves 13 is two. The two limiting grooves 13 are symmetrically distributed, and the two sides of the positioning groove are open. The inner walls of the positioning groove are provided with groove inner inclined surfaces 14 on both sides. The groove inner inclined surfaces 14 on both sides are symmetrical. The special-shaped bolt 3 includes a bolt column 31, and a special-shaped head 32 is fixedly connected to the bolt column 31. The side of the special-shaped head 32 is provided with an inclined surface section 33. The inclined surface section 33 is provided corresponding to the groove inner inclined surface 14, and the special-shaped bolt 3 is provided corresponding to the limiting groove 13. Specifically, a gasket 5 is sleeved on the outer side of the bolt column 31 , the gasket 5 contacts the outer side of the turbine casing 2 , and a self-locking nut 6 is threadedly connected to the bolt column 31 .

[0036] Furthermore, by cooperating with the inclined surface of the limiting groove 13 and the special-shaped bolt 3, firstly, the inclined surface matching structure only allows tangential sliding when the temperature changes, eliminating the normal displacement tendency, and ensuring that the contact surface always maintains a close fit; Secondly, the symmetrically distributed inclined surface design enables the thermal expansion force to be balanced in both directions, avoiding local stress concentration. The precise fit between the inclined surface section 33 of the special-shaped bolt 3 and the inclined surface 14 in the groove not only provides reliable circumferential constraints, but also can adapt to thermal deformation, thereby improving the positioning stability of the turbine outer ring 1 under high-temperature conditions, while ensuring the coordination of thermal deformation and solving the thermal stress problem that is easily caused by traditional rigid connections.

[0037] Specifically, one end of the spring 4 contacts the limiting groove 13 , and the other end of the spring 4 contacts the special-shaped head 32 , and the spring 4 is in a compressed state.

[0038] It should be noted that, since the working environment is at a high temperature, the spring 4 provided between the special-shaped bolt 3 and the limiting groove 13 needs to be made of a special high-temperature resistant material to ensure the stability of the structure. Furthermore, the compressed spring 4 continuously provides preload, ensuring that the inclined surface matching structure always maintains close contact under vibration conditions, suppressing high-frequency fretting wear. The spring 4 made of high-temperature resistant material can still maintain stable elastic properties in extreme temperature environments, avoiding connection failure caused by thermal relaxation in traditional structures. Secondly, the buffering effect of the spring 4 can absorb the instantaneous vibration energy of the turbine outer ring 1 and reduce the stress peak; at the same time, the elastic preload automatically compensates for the gap changes caused by thermal deformation, achieving continuous and stable connection under dynamic working conditions.

[0039] When the engine is in operation, the thermal expansion difference between the turbine casing 2 and the turbine outer ring 1 is relatively large. The scaling of the spring 4 can solve the problem of thermal expansion mismatch between the ceramic-based composite material and the metal material, ensuring that the turbine outer ring 1 and the turbine casing 2 always maintain a close fit under any operating condition of the engine, reducing the vibration of the turbine outer ring 1 under mechanical external force and aerodynamic load, and making the tip clearance 8 within a controllable range under any operating condition of the engine, thereby avoiding the turbine rotor 7 from rubbing against each other during high-speed rotation and damaging the high-speed rotating turbine blades, causing major failures.

[0040] In summary, the turbine outer ring 1 is located between the engine turbine casing 2 and the turbine blades. It is a typical hot end stator component. Its function is to reduce the leakage of the blade tip clearance and protect the engine casing from erosion by high-temperature leaking gas. The turbine outer ring 1 is made of a ceramic-based composite material, while the turbine casing 2, spring 4, bolts, and self-locking nut 6 are made of a high-temperature alloy. A beveled retaining groove 13 is provided on the turbine outer ring 1. The retaining groove 13 and the matching beveled surface of the special-shaped bolt 3 are pressed against each other before being tightened with the self-locking nut 6, thereby securing the composite outer ring to the metal turbine casing 2. Due to the significant difference in linear expansion coefficients between the SiCf / SiC composite material and the high-temperature alloy connector, thermal expansion differences occur between the turbine outer ring 1 and the turbine casing 2 under different engine operating conditions. The beveled fit of the groove on the turbine outer ring 1 ensures that relative sliding only occurs tangentially along the contact bevel during temperature changes, eliminating any tendency for normal displacement. The contraction of the spring 4 mitigates thermal mismatch, ensuring a constant close fit between the contact transfer surfaces of the two components, enabling a reliable radial mechanical connection between components with different thermal expansion coefficients. The expansion and contraction of the spring 4 ensures a constant close fit between the turbine outer ring 1 and the metal turbine casing 2 under all engine operating conditions, preventing friction during high-speed rotation of the turbine rotor 7.

[0041] The above embodiments are only intended to help understand the method and core concept of the present invention. It should be noted that, without departing from the principles of the present invention, a number of improvements and modifications may be made to the present invention by those skilled in the art, and such improvements and modifications also fall within the scope of protection of the claims of the present invention.

Claims

1. An embedded ceramic matrix composite material turbine outer ring for an aircraft engine, comprising a turbine casing (2), characterized in that: A turbine outer ring (1) is provided inside the turbine casing (2), and the number of the turbine outer rings (1) is multiple, and the multiple turbine outer rings (1) are distributed in a circumferential manner. A special-shaped bolt (3) is provided between the turbine outer ring (1) and the turbine casing (2), and the special-shaped bolt (3) is provided corresponding to the turbine outer ring (1). The special-shaped bolt (3) passes through the turbine casing (2), and a spring (4) is provided inside the turbine outer ring (1). A hollow area is formed on the turbine casing (2), and a turbine rotor (7) is provided on the turbine casing (2). The turbine rotor (7) is located in the hollow area, and a blade tip gap (8) is provided between the turbine rotor (7) and the turbine casing (2).

2. The embedded ceramic matrix composite turbine outer ring for an aircraft engine according to claim 1, characterized in that: The turbine casing (2) is provided with an inner annular surface (21), an annular groove (22) is provided on the inner annular surface (21), and the annular groove (22) is provided corresponding to the turbine outer ring (1).

3. The embedded ceramic matrix composite turbine outer ring for an aircraft engine according to claim 2, characterized in that: A positioning hole (23) is provided on the turbine casing (2), and the positioning hole (23) is located in the annular groove (22). There are multiple positioning holes (23), and the multiple positioning holes (23) are distributed in a circular pattern.

4. The embedded ceramic matrix composite turbine outer ring for an aircraft engine according to claim 1, characterized in that: An outer ring gap (15) is provided between adjacent turbine outer rings (1), and the outer ring gap (15) is used to provide expansion space.

5. The embedded ceramic matrix composite turbine outer ring for an aircraft engine according to claim 1, characterized in that: Positioning plates (11) are provided on both sides of the turbine outer ring (1), and a positioning block (12) is provided on the turbine outer ring (1), wherein the positioning block (12) and the positioning plate (11) are arranged in an I-shape.

6. The embedded ceramic matrix composite turbine outer ring for an aircraft engine according to claim 5, characterized in that: The positioning block (12) is provided with a limiting slot (13), the number of the limiting slots (13) is two, and the two limiting slots (13) are symmetrically distributed.

7. The embedded ceramic matrix composite turbine outer ring for an aircraft engine according to claim 6, characterized in that: Both sides of the positioning groove are open, and inner groove inclined surfaces (14) are provided on the inner walls of both sides of the positioning groove, and the inner groove inclined surfaces (14) on both sides are symmetrical.

8. The embedded ceramic matrix composite turbine outer ring for an aircraft engine according to claim 1, characterized in that: The special-shaped bolt (3) comprises a bolt column (31), a special-shaped head (32) is fixedly connected to the bolt column (31), a side of the special-shaped head (32) is provided with an inclined surface section (33), the inclined surface section (33) is provided corresponding to the inclined surface (14) in the groove, and the special-shaped bolt (3) is provided corresponding to the limiting groove (13).

9. The embedded ceramic matrix composite turbine outer ring for an aircraft engine according to claim 1, characterized in that: One end of the spring (4) contacts the limiting groove (13), and the other end of the spring (4) contacts the special-shaped head (32), and the spring (4) is in a compressed state.

10. The embedded ceramic matrix composite turbine outer ring for an aircraft engine according to claim 8, characterized in that: A gasket (5) is sleeved on the outer side of the bolt column (31), and the gasket (5) contacts the outer side of the turbine casing (2). A self-locking nut (6) is threadedly connected to the bolt column (31).

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